Gas turbine engines include machinery that extracts work from combustion gases flowing at high temperatures, pressures and velocity. The extracted work can be used to drive a generator for power generation or for providing the required thrust for an aircraft. A typical gas turbine engine consists of a multistage compressor where the atmospheric air is compressed to high pressures. The compressed air is then mixed at a specified fuel-air ratio in a combustor wherein its temperature is increased. The high temperature and pressure combustion gases are then expanded through a turbine to extract work so as to provide the required thrust or drive a generator or a compression device depending on the application. The turbine includes at least a single stage with each stage consisting of a row of blades and a row of vanes. The blades are circumferentially distributed on a rotating hub with the height of each blade covering the hot gas flow path. Each stage of non-rotating vanes is placed circumferentially, which also extends across the hot gas flow path. The included invention involves the combustor and turbine sections of gas turbine engines, each of which will be further discussed.
The combustor portion of a gas turbine engine can be of several different types: silo, can/tubular, annular, and a combination of the last two forming a can-annular combustor. It is in this component that the compressed fuel-air mixture passes through fuel-air nozzles and a combustion reaction of the mixture takes place, creating a hot gas flow causing it to drop in density and accelerate downstream. The can combustor typically comprises of individual, circumferentially spaced cans that contain the flame of each nozzle separately. Flow from each can is then directed through a duct and combined in an annular transition piece before it enters the first stage vane. In the annular combustor type, fuel-air nozzles are typically distributed circumferentially and introduce the mixture into a single annular chamber where combustion takes place. Flow simply exits the downstream end of the annulus into the first stage turbine, without the need for a transition piece. The key difference of the last type, a can-annular combustor, is that it has individual cans encompassed by an annular casing that contains the air being fed into each can. Each variation has its benefits and disadvantages, depending on the application.
In combustors for gas turbines, it is typical to premix the fuel and the air before it enters the combustion chamber through a set of fuel-air nozzles. These nozzles introduce a swirl to the mixture for several reasons. One is to enhance mixing and thus combustion, another reason is that adding swirl stabilizes the flame to prevent flame blow out and it allows for leaner fuel-air mixtures for reduced emissions. A fuel air nozzle can take on different configurations such as single to multiple annular inlets with swirling vanes on each one.
As with other gas turbine components, implementation of cooling methods to prevent melting of the combustor material is needed. A typical method for cooling the combustor is effusion cooling, implemented by surrounding the combustion liner with an additional, offset liner, which between the two, compressor discharge air passes through and enters the hot gas flow path through dilution holes and cooling passages. This technique removes heat from the component as well as forms a thin boundary layer film of cool air between the liner and the combusting gases, preventing heat transfer to the liner. The dilution holes serve two purposes depending on its axial position on the liner: a dilution hole closer to the fuel-air nozzles will cool the liner and aid in the mixing of the gases to enhance combustion as well as provide unburned air for combustion, second, a hole that is placed closer to the turbine will cool the hot gas flow and can be designed to manipulate the combustor outlet temperature profile.
The next portion of a gas turbine engine that the flow travels through is the first stage vane and turbine. At this point in a gas turbine engine, the hot gases are further accelerated as well as turned to a velocity that allows it to strike a row of turbine blades that extract work from the hot gases by producing lift on the turbine blades which results in the rotation of a drive shaft. In such an application, the turbine blades and vanes in the hot gas path operate under conditions of high temperature, pressure, and velocity. These hostile conditions cause thermal oxidation and surface deterioration leading to reduced component life. Inlet turbine gas temperatures typically reach about 200-300° C. above the melting point of turbine components. These high temperatures significantly deteriorate the surface conditions and increase the surface roughness; therefore, it is important that these surfaces be cooled. A variety of designs, materials and configurations are used in gas turbine engines that provide structural robustness as well as effectively cool the turbine vanes and blades in order to enhance its durability against hot combustion gases; however, there has been no attempt at modifying the combustor and turbine in such a way as to eliminate the need for the first row vanes entirely. Currently, this first row of turbine vanes require the development of various technologies in order to cope with the extreme operating environment that include but are not limited to: expensive nickel-alloys, thermal barrier coating, complex casting methods to incorporate internal cooling passages, and filming cooling techniques. In some cases the first row vanes can represent approximately 5% of the complete gas turbine engine cost. In addition, approximately 2% of the total flow losses through a gas turbine can be attributed to pumping cooling air through this single component. This invention will function in a manner consistent with today's gas turbines; however, it will do so without the first stage vane nozzle thus eliminating the associated issues of cost and performance losses.